hypersonic aerody namic appraisal of winged blunt, rather sharp and spatuled body … ·...

1

Abstract

The paper deals with the hypersonic

aerodynamic analysis of three reusable and

unmanned flying laboratories designed to

perform a return flight from low Earth orbit to

provide experimental data in the framework of

re-entry technologies. Several design

approaches, ranging from low-order methods to

computational fluid dynamics analyses, have

been addressed in this work. In particular,

vehicles aerodynamic performances for a wide

range of free stream flow conditions, including

reacting and non-reacting flow and different

angles of attack, have been provided and in

some cases compared. Computational fluid

dynamics results confirm that real gas effects

seem to be fundamental for the assessment of

the concept aerodynamics, especially

concerning pitching moment evaluation.

1 Introduction

This paper deals with the aerodynamic

performance analysis of three reusable and

unmanned flying laboratories designed to

perform an experimental flight return from low

Earth orbit. Therefore, each vehicle concept

belongs to the class of orbital re-entry vehicle

(ORV) e.g., re-entry energy of the order of 25

MJ/kg. Indeed, concepts under investigation in

the present research effort are conceived as

flying test beds (FTB) that will re-enter the

Earth’s atmosphere, thus allowing to perform a

number of experiments on critical re-entry

technologies. For example, the FTB may be

useful to demonstrate maneuverability in the

upper atmosphere, to test advanced thermo-

structure concepts, such as leading edges made

of advanced thermal protection material (TPM),

and to investigate the flowfield features during

re-entry in order to validate numerical (e.g.,

CFD) and experimental prediction capabilities.

In particular, the vehicle may provide

aerodynamic and aerothermodynamic flight data

to correlate with ground test (e.g., the CIRA

Plasma Wind Tunnel “Scirocco”) results, thus

providing new insights into the understanding of

complex aerothermodynamic phenomena

occurring in flight and improving prediction

methodologies and extrapolation to flight

capabilities.

Right now Europe has undertaken the

development of three very different FTBs,

namely ARD (atmospheric re-entry

demonstrator), Expert (european experimental

reentry testbed), and IXV (intermediate

experimental vehicle). ARD was a scaled-down

version of an Apollo capsule. It was launched

by ARIANE 5 V503 on October 21, 1998. After

a fully successful sub-orbital and re-entry flight,

it was recovered in the Pacific Ocean [1]. ARD

allowed Europe to assess the aerodynamics of

such a kind of capsule that still represents a very

attractive design solution for what concerns

manned high energy re-entry (e.g., return from

Mars/Moon missions). Expert, not yet flown, is

a small sphere-cone FTB designed to perform

several in-flight experiments, such as for

example advanced thermal protection system

(TPS), wall catalyticity, flow transition

assessment and so on [2]. Finally, the

Intermediate Experimental Vehicle, which is

still under development by the European Space

Agency (ESA), is a rather blunt FTB which

HYPERSONIC AERODYNAMIC APPRAISAL OF WINGED BLUNT, RATHER SHARP AND SPATULED

BODY RE-ENTRY VEHICLES

Antonio Viviani*, Giuseppe Pezzella**

* Second University of Naples, via Roma, Aversa. Italy,

** Italian Aerospace Research Centre, via Maiorise, Capua. Italy

ANTONIO VIVIANI, GIUSEPPE PEZZELLA

2

features a lifting-body configuration. It will face

re-entry flight conditions in the fall 2014 at the

end of a sub-orbital flight characterized by an

energy level very close to that of an orbital re-

entry (e.g., 25 MJ/kg) [3]. IXV will allow to

address several in-flight experiments like

GN&C of a flapped aeroshape, TPS catalyticity,

and etc. The aerodynamic characterization of

IXV can be found in [4]. Generally speaking, a

reusable ORV operates at different flight

regimes from subsonic to hypersonic speeds. A

typical mission profile includes: ascent phase,

where the spacecraft is attached to a launch

vehicle and placed at an altitude; orbit phase,

where vehicle orbits in space till completion of

desired mission; and descent phase, where the

ORV re-enters in atmosphere and lands like a

conventional airplane for subsequent use.

During the descent up to landing phase the

spacecraft encounters subsonic speeds.

Therefore, the choice of vehicle aeroshape and

its aerodynamic characterization at hypersonic

speeds is of vital importance for safe

return.Usually, the vehicle configuration is

continuously adapted throughout the design

phase by means of a multidisciplinary trade-off

study involving several concepts. Of course the

winning configuration from the aerothermal

point of view is the one showing, at the same

time, the best aerodynamic and

aerothermodynamic performances. Right now

the most promising vehicle configurations,

resulted from trade-off design analyses, are

shown from Figure 1 to Figure 4.

Figure 1 Rather blunt (up), sharp (middle) and spatuled

body configurations.

Figure 2 shows the rather sharp vehicle

configuration, namely ORV-WSB. In this figure

the concept appears also docked with the service

module with deployed solar panels (e.g., orbital

stationing phase) [5][6].

Figure 2 Rather sharp configuration, namely ORV-WSB

and the service module with solar panels [5][6].

The rather blunt configuration, named ORV-

WBB is provided in Figure 3 [6].

Figure 3 Rather blunt vehicle configuration, named

ORV-WBB [6].

Finally, Figure 4 displays a spatuled-body (SB)

configuration, namely ORV_SB [6].

Figure 4. The Spatular body configuration, namely ORV-

SB [6].

It is worth to note that ORV concepts show

different aeroshapes to address different

experimental investigation aims. For example,

the ORV-SB configuration is most attractive

considering that it represents the only viable

way to accomplish and optimize the integration

3

HYPERSONIC AERODYNAMIC APPRAISAL OF WINGED BLUNT, RATHER SHARP AND SPATULED BODY RE-ENTRY VEHICLES

of scramjet propulsion with the vehicle

aerodynamic configuration (see Figure 5, where

the ORV-SB features a scramjet engine on the

belly side), thus evolving toward waverider

aeroshape [7].

Figure 5. The Spatular body configuration with scramjet

engine.

Indeed, since the beginning of aviation, the

trend in aircraft design has been towards greater

speed. The next frontier of speed envelope is

travel at hypersonic speeds. One of the most

practical and efficient approach to travel at these

high speeds is known as the waverider. Figure 6

shows that concerning high-performance flight

vehicle architecture converges with the

technology of airbreathing configuration.

Figure 6 Space and atmospheric vehicle development

coverage, so the technology of high-performance

launchers converges with the technology of airbreathing

aircraft. M=Mach number [7]

Such a configuration demands high

aerodynamic efficiency [7]. Indeed, the most

efficient hypersonic lifting surface is the

infinitely thin flat plate, provided that its lift-to-

drag ratio is the highest that can be achieved at

hypersonic speeds. The flat plate, however, is

obviously not practical, especially since it

cannot contain any volume for payload, engines,

fuel, etc. Therefore, a more realistic

configuration design converges to a spatular

vehicle architecture. The characteristics of this

aeroshape are: very small frontal area and

highly streamlined configuration to minimize

total surface area; very little wing area, but the

fuselage is often shaped to generate additional

lift; and propulsion assembly highly integrated

into the vehicle fuselage.

Anyway, all concepts under investigation

belong to the class of the winged body vehicles.

Such configurations, however, differ in terms of

several vehicle’s features as for example

planform shape, cross section, nose camber,

wing swept angle and, vertical empennages.

Differences in concept aeroshapes can be clearly

appreciated in Figure 7, where each aeroshape is

over imposed on each other.

Figure 7 ORV aeroshapes comparison

In this framework, this research effort provides

an overview of the aerodynamic performances

of these ORVs at hypersonic speed in

continuum flow condition. Both low-order

methods (i.e., hypersonic panel methods) and

CFD design analysis have been considered to

assess vehicle aerodynamic characteristics,

compliant with a phase-A design level. Low-

order methods design approach has been

extensively used; while CFD simulations are

performed to address the reliability of low-order

method design results and to investigate on

complex flowfield phenomena not predictable

with simplified tools [8],[9],[10]. Indeed, the

range between Mach 2 to Mach 25 is analyzed

and both perfect and reacting gas CFD

simulations are performed at several points of

the flight scenario. At flight conditions where

real gas effects occur the air is modeled as a

mixture of five species in thermo-chemical non-

equilibrium conditions. In fact, it is well known

that the pitching moment can be highly

modified by high temperature effects, thus

affecting vehicle’s stability behavior and

trimming conditions [11]. Finally, an analysis of


4

the longitudinal and lateral-directional stability

has been also provided for one concept, together

with some of main interesting features of the

flowfield past the vehicles at different Mach

numbers.

2 Vehicles Description and Flight Scenario

Vehicle concepts feature a compact wing-body

configuration equipped with a rounded edge

delta-like fuselage cross section, a delta wing,

and V-tail. The vehicle architecture shows a

blended wing body interface and a flat bottomed

surface to increase its overall hypersonic

performance. The fuselage was designed to be

longitudinally tapered, in order to improve

aerodynamics and lateral-directional stability,

and with a cross section large enough to

accommodate all the vehicle subsystems. The

last fuselage’s feature has a large impact on

vehicle performance. In fact, from the

aerodynamic point of view, the lift and the

aerodynamic efficiency are mainly determined

by the fuselage fineness and by the shape of the

vehicle cross section [12]. The forebody is

characterized by a rather simple cone-sphere

geometry with smooth streamlined surfaces on

the upper and lower side of fuselage, and by the

nose drop-down configuration, typical of

winged hypersonic vehicles. The nose camber is

low enough to reduce elevons size in order to

provide desired trim range and to improve

internal packaging. The wing size and location

were defined on the basis of trade-off studies so

to improve vehicle aerodynamics and to provide

static stability and controllability during flight

[8],[10],[13].

Finally, the wing is swept back to assure

best performance with respect to supersonic

drag and aerodynamic heating. A properly

designed strake could be added in the future,

depending on the confirmation of a specific

landing requirement. A wing dihedral angle of 5

deg is also provided to enhance vehicle lateral-

directional stability. The wing also features a

high length-to-width ratio to minimize drag, and

a section shape that is maintained from root to

wing tip; a leading edge that is rather sharp and

a nearly flat bottomed surface to dissipate

efficiently the aeroheating. Vertical tails sweep

angle is 45 deg. Control power for vehicle is

provided by two wing-mounted elevon surfaces

(which must serve as ailerons and elevators),

and rudders surface. Used symmetrically

elevons are the primary controls for the pitch

axis, i.e., pitch control. Roll control is obtained

through asymmetrical usage of these elevons.

Rudders help to provide the directional

control, i.e., sideslip stability. During entry,

when the vehicle is flying at high angles of

attack, rudders should be augmented by reaction

control system (RCS). Then, the vehicle may be

provided by a body flap located at the trailing

edge of the fuselage in order to augment pitch

control and to shield the nozzle of propulsion

subsystem (ORV-WBB only). Indeed, trim

capability to relieve elevon loads is obtained by

body flap deflection. At hypersonic speeds a

surface behind the vehicle CoG balances the

nose up pitching moment typical of such kind of

configuration at hypersonic speeds.

Finally, the aerodynamic control surfaces

are large enough to provide stability without

sacrificing too much lift.

In the framework of the Flight Mechanics

trade-off analyses several re-entry trajectories

have been computed thus defining a vehicle

flight envelope. For example, Figure 8 shows a

number of re-entry trajectories evaluated

supposing that the vehicle heat flux constraint

ranges between 1.1 and 2 MW/m2; while the

dynamic pressure limit is equal to 12 kPa. The

Mach-Reynolds numbers grid in the altitude-

velocity map of Figure 8 is also reported in

order to characterize vehicle aerodynamics,

according to the space-based design approach

[14]. USV-3 Flight Envelope

0

20000

40000

60000

80000

100000

120000

0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Velocity , m/s

Alt

itu

de

, m

M∞=5 M

∞=10 M

∞=16 M

∞=20 M

∞=25

Re∞Lref=1x104

Re∞Lref=1x105

Re∞Lref=5x10

5

Re∞Lref=1x10

6

Figure 8. The ORV flight envelope in the velocity-altitude

map.

5


3 Design Approach and Used Tools

A summary overview of the aerodynamic

performances of vehicle concepts, compliant

with a phase-A design level, is herein performed

[13], [14], [15]. The activities goal is to provide

concept aerodynamic database (AEDB) for

Flight Mechanics and thermal shield design

analyses. In fact, it must be verified that vehicle

is able to stay within the load constraints (i.e.,

re-entry corridor) during descent up to

conventional runway landing.

The concepts are characterized by a number of

extreme loading flight conditions as, for

example, transitional flow regime and peak

heating conditions. It must return from orbit, fly

trimmed throughout hypersonic and supersonic

regimes until landing is gained. The AEDB has

been provided as a function of Mach number,

angle of attack, sideslip angle, aerodynamic

control surface deflections, and Reynolds

number, according to the Space-Based design

approach [14]. This design approach dictates the

generation of a complete dataset as function of a

number of independent parameters (i.e. M∞,

Re∞, α, β) as schematized in Figure 9.

H

V

H

V

H

V

Figure 9 Space-based design approach in the altitude-

velocity map.

An accurate aerodynamic analysis of all these

flight conditions, however, is very complex and

time consuming, and is not compatible with a

Phase-A design study, in which fast predicting

methods are mandatory. Therefore, the

evaluation of the vehicle AEDB was mainly

performed by means of engineering tools, while

a limited number of more reliable CFD

computations (continuum regime only) were

performed in order to verify the attained

accuracy and to focus on some critical design

aspects not predictable with simplified tools

[15], [16], [17]. This overall process is referred

to as “anchoring” of the engineering level

methods. The anchoring process permits a few,

select, CFD solutions to be used beyond the

specific flight conditions at which they were

original run. Moreover, the anchoring process

allows for the cost effective use of high fidelity,

and computationally expensive, CFD solutions,

early in the design process when the vehicle

trajectories are often in a constant state of

change [16], [17]. The CFD anchoring “space”

is defined by a small number of CFD solutions

in Reynolds-Mach-AoA space as shown in

Figure 10.

CFD # n-1

Mach

AoA

Reynolds number

CFD # n

CFD #1

CFD #2

Engineering result

CFD result

[1÷70] x106

[0÷50] deg

[2÷20]

CFD # n-1

Mach

AoA

Reynolds number

CFD # n

CFD #1

CFD #2

Engineering result

CFD result

Engineering result

CFD result

[1÷70] x106

[0÷50] deg

[2÷20]

Figure 10 Hypothetical CFD anchoring mesh in

Reynolds-Mach-AoA space.

In the framework of low-order methods codes

vehicle aerodynamics has been addressed by

means of HPM code; while CFD analysis, both

Euler and Navier-Stokes, for hypersonic speed

have been carried out with commercial code

FLUENT.

The HPM code is a 3-D Supersonic-

Hypersonic Panel Method code, developed at

CIRA, that computes the aerodynamic

characteristics of complex arbitrary three-

dimensional shapes by using surface inclination

methods (SIM) typical of Newtonian

aerodynamics [11], [13], [18]. Detailed code

description and features can be found in [19],

[20], [21].

The CFD code FLUENT solves the full

Reynolds Averaged Navier-Stokes equations in

a finite volume approach, with a cell centred

formulation on a multi-zone block-structured

grid. In the present research effort the thermal

and chemical non-equilibrium flowfield

governing equations are integrated in a density-

based approach with an upwind Flux Difference

Splitting (FDS) second order upwind numerical

scheme for the spatial reconstruction of the


6

convective terms; while for the diffusive fluxes

a cell-centred scheme is applied. In some

computations, however, flux vector was

computed by using a flux-vector splitting

scheme, namely Advection Upstream Splitting

Method (AUSM). It provides exact resolution of

contact and shock discontinuities and it is less

susceptible to Carbuncle phenomena. Implicit

solver formulation was considered in the

computations of this work. Indeed, due to

broader stability characteristics of the implicit

formulation, a converged steady-state solution

can be obtained much faster using the implicit

formulation rather than the explicit one. Global

transport properties of the gas mixture relied on

semi-empirical rules such as Wilke’s mixing

rule for viscosity and thermal conductivity. The

viscosity and thermal conductivity of ith

species

was obtained by kinetic theory of gases [11].

For the diffusion coefficient of the ith

species in

the mixture the multi-component diffusion

coefficient was applied, where species mass

diffusivity is evaluated by kinetic theory [11].

Flowfield chemical reactions proceed with

forward rates that are expressed in the Arrhenius

and reaction rate parameters are due to Park

[22]. In particular, a number of in-house

modifications (i.e., User Defined Functions –

UDF–) for the thermal non-equilibrium were

considered since vibrational non-equilibrium

conditions are not basic code features. In the

UFD vibrational relaxation is modeled using a

Landau-Teller formulation, where relaxation

times are obtained from Millikan and White,

assuming simple harmonic oscillators [11].

Finally, the k-ω SST model has been considered

to account for turbulence effects and only steady

state computations are carried out so far.

4 Aerodynamic Characterization

Concepts aerodynamic force and moment have

been non-dimensionalized by means of the

following reference parameters: Lref=1.26 m

(wing mean aerodynamic chord–longitudinal

reference length); cref=3.60 m (lateral-

directional reference length); Sref=5.18 m2

(reference Surface).

4.1 Low order methods aerodynamics

Simplified aerodynamic analysis for supersonic

and hypersonic speeds was accomplished on

panel mesh close to that shown in Figure 11.

Figure 11 One of the ORV-WSB panel mesh.

In this figure is reported, for example, one of the

surface panel mesh considered for the ORV-

WSB aerodynamic assessment. Trade-off design

analyses highlighted that the best surface

inclination methods to consider in assessing

vehicle aerodynamic performance are tangent

cone and tangent wedge empirical methods for

fuselage and wing belly side, respectively; while

newtonian method (i.e., Cp=0) at vehicle leeside

[23],[24]. Some of main results obtained for the

ORV-WSB in clean configuration aerodynamic

(i.e., no aerodynamic surface deflected) are

shown from Figure 12 to Figure 16 .

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.50

0.05 0.10 0.15 0.20 0.25

Drag Coefficient

Lift C

oeffic

ient

M=2

M=3

M=4

M=5

M=6

M=7

M=8

M=9

Mach

Figure 12 ORV-WSB aerodynamic polars for 2≤M∞≤9

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6

Drag Coefficient

Lift C

oeffic

ient

M=10

M=16

M=25

Figure 13 ORV-WSB aerodynamic polars for 10≤M∞≤25

7


pole @ CoG

-0.7

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0.0

0.1

0.2

0 5 10 15 20 25 30 35 40 45

Angle of Attack, (deg)

Pitchin

g m

om

ent co

effic

ient

M=2

M=3

M=6

M=9

Figure 14 ORV-WSB CMy for 2≤M∞≤9

pole @ CoG

-0.30

-0.20

-0.10

0.00

0.10

0.20

0 5 10 15 20 25 30 35 40 45

Angle of Attack, (deg)

Pitch

ing

mo

men

t coeffic

ient

M=10

M=16

M=25

Figure 15 ORV-WSB CMy for 10≤M∞≤25

Figure 14 shows the aerodynamic polars and the

pitching moment coefficients for Mach ranging

from 2 to 25 and α from 0 to 40 deg. As shown,

ORV-WSB drag and lift decrease as Mach

number increases up to reach a value that does

not change even if Mach still rises, according to

the Oswatich principle [11],[13]. Figure 14 also

shows that the configuration is statically stable

(i.e., CMyα< 0) for α higher than 20 deg in

hypersonic conditions [13]. In particular, the

concept in clean configuration features a natural

trim point (i.e., CMy=0) at about 33 to 38 deg

AoA for M∞=6 and 9, respectively. At higher

Mach number trim AoA ranges from about 31

to 33 deg. Moreover, the pitching moment

behavior highlights that ORV-WSB can be

trimmed through flap positive deflection (i.e.,

downward) only at hypersonic speed. At M∞=2

and M∞=3 pitching moment derivative is

negative for α larger than 5 deg and 15 deg,

respectively. As far as the lateral-directional

stability is concerned, Figure 16 shows for α=5

deg the effect of sideslip on rolling (CL) and

yawing moment (CN) coefficients along with

Mach number. Recall that for directional

stability, CNβ>0 and for dihedral effect, CLβ<0

[13].

-0.0006

-0.0004

-0.0002

0.0000

0.0002

0.0004

0.0006

0.0008

0.0010

0.0012

0 1 2 3 4 5 6 7 8 9 10

Mach number

CLb

CNb

Figure 16 ORV-WSB effect of sideslip on CLβ and

CNβ up to M∞=9, at α=5 deg.

As shown, the configuration is statically stable

in lateral-directional flight at α=5 deg. Note

that, the body flap can obviously offer

advantages also on both longitudinal and lateral-

directional stability by providing margins on

CoG location. In fact, the body flap, located on

the rear lower portion of the aft fuselage, allows

to pitch trim the vehicle while elevons providing

concept roll control.

Finally, aerodynamic efficiency and pitching

moment comparison at M∞=10 among ORV-

WSB, ORV-WBB, and ORV-SB can be

recognized in Figure 17 and Figure 18,

respectively.

-1.0

-0.5

0.0

0.5

1.0

1.5

2.0

2.5

3.0

0 10 20 30 40 50

L/D

AoA, (deg)

L/D at M=10ORV-SB

ORV-WSB

ORV-WBB

Figure 17. Lift-to-Drag ratio at Mach 10. Comparison

among concept performances.

As shown, the ORV-SB concept features the

best lift-to-drag ratio up to α=20 deg. In

particular, the (L/D)max is equal to about 2.8 and

is attained at α=10 deg AoA. On the other hand,


8

the maximum aerodynamic efficiency of ORV-

WSB and ORV-WBB is reached at about α=15

deg and is equal to about 2.4 and 1.8,

respectively. For AoA larger than 20 deg

differences in aerodynamic efficiency decrease

as α increases, and they vanish for α>35 deg.

-1.4

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

0 10 20 30 40 50 60

CM

y

AoA, (deg)

CMy at M∞∞∞∞=10 and pole at CoG

ORV-SB CoG at 63%

ORV-SB CoG at 56%

ORV-WSB CoG at 63%

ORV-WSB CoG at 56%

ORV-WBB CoG at 63%

ORV-WBB CoG at 56%

Figure 18. Effect of CoG position on CMy at Mach 10.

Comparison among concept performances

As a result, in the framework of re-entry at high

angle of attack, namely 35-40 deg (e.g., close to

that of US Orbiter) differences in aeroshape do

not significantly affect the descent flight. In

fact, at hypersonic speed and at high AoA

vehicle aerodynamic is dictated essentially by

its planform shape. In this case, ORV planform

shapes are quite close each other (see Figure 7).

As far as pitching moment is concerned, the

effect of CoG position with respect to the

fuselage length on the CMy for each vehicle

concept is summarized in Figure 18. As shown,

when the CoG is at 63% the ORV-SB concept

features a strong static instability in longitudinal

flight highlighting that the centre of pressure is

well ahead of the CoG (i.e., negative static

margin); ORV-WBB is static stable in pitch for

α>40 deg and can be trimmed by positive flap

deflections; while ORV-WSB is static stable in

longitudinal flight for α>30 deg and shows also

a natural trim point at about 45 deg AoA. On the

contrary for CoG at 56% the ORV-SB concept

becomes statically stable in longitudinal flight

for α>35 deg and trim angles of attack can be

attained by positive flap deflections; the other

two concepts (WB and WBB) are statically

stable in pitch for α>5 deg and feature a natural

trim point at 10 and 20 deg AoA, respectively.

In particular, they can be trimmed at high AoA

by means of negative (i.e., trailing edge up) flap

deflections. Anyway, pitching moment behavior

versus AoA points out that vehicle subsystems

arrangement (i.e., CoG position) must be

carefully addressed for a static stable and

trimmable vehicle concept.

4.2 CFD-based aerodynamics

For what concerns numerical flowfield

computations, on the base of the flight envelope

of Figure 8 a number of flight conditions has

been chosen to perform some CFD

computations in steady state conditions.

Numerical results aim to anchor engineering

analyses in order to increase them accuracy, and

to focus on some critical design aspects not

predictable by using simplified tools as, for

example, Shock-Shock Interaction (SSI)

phenomena on leading edges of both wing and

tail, and real gas effects as well [11], [12], [13].

The CFD test matrix is summarized in Table 1.

Mach 0 5 10 20 30 40 45 2 4 8

2 X X X X X

3 X X X X X X X

4 X X

5 X X X X X

6 X X X X

7 X X X X

8 X X

8 X X

10 X X X

16 X X X

20 X X X X

20 X X X X

25 X X

X Perfect Gas

X Reacting Gas

CFD test matrix

AoA @ AoS=0 deg AoS @ AoA=5 deg

Table 1 CFD test matrix.

Note that each cell identifies a CFD run (i.e.,

check point). Therefore, the results of forty

seven one turbulent CFD computations were

performed and some results reported

hereinafter. It is worth nothing that, at M∞=8,

10, 16, 20, and 25 non-equilibrium CFD

computations are also carried out. Real gas

effects can be important because, during

atmospheric re-entry, dissociation process takes

place in the shock layer, which can have an

influence on the aerodynamic coefficients. Real

gas effects are expected to influence stability

and control derivatives of vehicle, in particular

its pitching moment, as highlighted by first

Shuttle re-entry (STS-1) where an unexpected

higher nose-up pitching moment required a

9


body-flap deflection twice than that predicted

by the pre-flight analyses to trim the Orbiter

[11],[13]. Further, real gas effects cause a shock

that lies closer to the vehicle with respect to the

position that would characterize a perfect gas

case (i.e., thin shock layer) [11],[13]. These

effects obviously occur only at high Mach

numbers [11], [12], [13]. CFD computations

have been carried out on both a multi-block and

hybrid unstructured grids similar to those shown

in Figure 19 and Figure 20.

Figure 19. Computational domains. Mesh on symmetry

plane and ORV-WSB vehicle.

Figure 20. Computational domains. Mesh on vehicle

surface and ORV-WBB vehicle.

Figure 21. Computational domains. Mesh on

vehicle surface and ORV-SB vehicle.

In these figure close-up views of 3-D sup-

hypersonic mesh on both symmetry plane and

vehicle surface can be seen in the case of ORV-

WSB (left) and ORV-WBB (right side). Of

course, for both computational domains the

distribution of surface grid points was dictated

by the level of resolution desired in various

areas of vehicles such as stagnation region,

according to the computational scopes.

Grid refinement in strong gradient regions of

flowfield was made through a solution adaptive

approach. The coordinate y+ of the first cell

adjacent to the surface is about 1.

As far as numerical results are concerned, it is

worth to note that they refer to both converged

and grid independent computations

[5],[6],[15],[19].

Indeed, aerodynamic results comparison

between numerical and engineering analysis in

the case of M∞=10, 20 and for 0<α<40 deg are

provided in Figure 22 and Figure 23.

M=10 Euler CFD

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

0 10 20 30 40 50

Angle of Attack, deg

Lif

t C

oeff

icie

nt

CFD

HPM

M=20 Euler CFD

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

0 10 20 30 40 50


Lift

Co

eff

icie

nt

CFD-PGCFD-RGHPM

Figure 22. CL vs α. HPM and CFD results comparison at

M∞=10 and 20. ORV-WSB concept.


10

M=10 Euler CFD

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

0 10 20 30 40 50


Dra

g C

oe

ffic

ien

t

CFD

HPM

M=20 Euler CFD

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

0 10 20 30 40 50


Dra

g C

oe

ffic

ien

t

CFD-PGCFD-RGHPM

Figure 23. CD vs α. HPM and CFD results comparison at

M∞=10 and 20. ORV-WSB concept.

As shown, results comparison confirms that

engineering-based (i.e., HPM) and numerical

data fit each other very well. Moreover,

differences between real gas and frozen gas

coefficients, collected also in Table 2 for

M∞=20 and 10<α<40 deg, suggest that high

temperature gas effects are negligible for lift

and drag; whereas a significant effect on vehicle

pitching moment is expected along the re-entry

trajectory varying from 8 to 50%.

AoA PG RG Err %

10 0.14 0.14 -1.15

20 0.49 0.48 1.37

30 0.93 0.91 1.97

40 1.33 1.31 1.46

AoA PG RG Err %

10 0.06 0.06 3.80

20 0.22 0.22 2.00

30 0.59 0.58 1.72

40 1.18 1.16 0.94

AoA PG RG Err %

10 0.036 0.043 -17.01

20 0.086 0.094 -7.98

30 0.082 0.096 -14.55

40 0.025 0.048 -47.82

CL

CD

CM

Table 2 Real Gas effects on ORV-WSB aerodynamics

Further, in the following figures some of main

interesting hypervelocity flowfield features

obtained for the ORV-WSB concept are shown.

For example, the Mach number and flow

species fields for M∞=10 and 20 at AoA=30,

and 40 deg are reported from Figure 24 to

Figure 25.

Looking at contour fields on the vehicle

symmetry plane, one can appreciate the rather

strong bow shock that occurs ahead of the

vehicle concepts during descent at these flight

conditions, especially at higher AoA. Indeed,

Figure 24 shows the Mach number field past the

rather sharp wing body vehicle for M∞=10 and

30 deg AoA [25]. As shown, even if the CFD

computations are carried out in the case of

perfect gas flow, the bow shock is very close to

the vehicle due to its streamlined configuration

Figure 24. Mach contours on symmetry plane and ORV-

WSB surface at M∞=10 and α=30 deg.

In the reacting gas computations air was

considered made up of five chemical species

(O2, N2, O, N and NO) and in thermo-chemical

non-equilibrium conditions.

For example, the Oxygen (O) mass fractions

contour fields at H∞=70 km, M∞=20 and α=40

deg are shown in Figure 25.

As far as numerical results for the ORV-WBB

concept are concerned, Figure 26 to Figure 28

show several flowfield features for the rather

blunt vehicle at M∞=25 and α=30 deg.

11


Figure 25. Oxygen mass fractions. CFD computation at

H∞=70 km, M∞=20 and α=40 deg. ORV-WSB concept.

In particular, Figure 26 and Figure 27 highlight

the Mach number and the N2 mass fraction

contour fields on the vehicle symmetry plane

and Cp on the concept outer surface,

respectively.

Figure 26. Mach contours on symmetry plane and Cp on

ORV-WBB surface at M∞=25 and α=30 deg..

Figure 27. N2 contours on symmetry plane and Cp on

ORV-WBB surface at M∞=25 and α=30 deg.

Side and bottom views of the surface

distribution of Cp at these flight conditions are

summarized in Figure 28.

Figure 28. Cp on ORV-WBB surface at M∞=25 and α=30

deg.

Hence, in order to get an idea of the bow shock

that envelopes the ORV-WBB at M∞=25 and

α=30 deg, Figure 29 provides three flowfield

cross sections where Mach contours are

reported.

Figure 29. Cp on ORV-WBB surface at M∞=25 and α=30

deg. Mach number contours an three flowfield cross

sections

For what concerns the flowfield past the ORV-

SB concept, Figure 30 shows the Mach number

contour field on vehicle symmetry plane and

pressure distribution on concept surface at

M∞=25 and α=30 deg. Streamlines are also

provided.


12

Figure 30. Pressure on ORV-SB surface and Mach

number contours on symmetry plane at M∞=25 and α=30

deg.

3-D streamtraces at these flight conditions

together with a cross flow section coloured by

Mach number are provided in Figure 31.

Also in this case a very narrow shock layer is

expected for this concept.

Figure 31. Pressure on ORV-SB surface at M∞=25 and

α=30 deg with 3-D streamtraces and flowfield cross

section coloured by Mach number.

Finally, results comparison between numerical

and engineering-based aerodynamic for ORV-

WBB and ORV-SB are summarized in Figure

32 and Figure 33, respectively.

For instance, Figure 32 compares aerodynamic

results for axial force coefficient (CA) and

pitching moment coefficient (CMy) at M∞=25 and

α=30 deg; while Figure 33 provides result

comparison for normal force coefficient (CN) at

M∞=10 and α=30 deg and for axial force

coefficient at M∞=25 and α=30 and 45 deg.

0,10

0,11

0,12

0,13

0,14

0,15

0,16

0,17

0,18

0,19

0,20

0 10 20 30 40 50

CA

AoA, deg

M=25

SIM

CFD

-14,0

-12,0

-10,0

-8,0

-6,0

-4,0

-2,0

0,0

2,0

0 10 20 30 40 50

CM

AoA, deg

M=25

SIM

CFD

Figure 32. CA and CM at M∞=25 and α=30 deg. ORV-

WBB concept. CFD and HPM comparison.

-0,5

0,0

0,5

1,0

1,5

2,0

2,5

3,0

3,5

4,0

0 10 20 30 40 50

AoA, deg

M=10

SIM

CFD

0,00

0,02

0,04

0,06

0,08

0,10

0,12

0,14

0,16

0 10 20 30 40 50

CA

AoA, deg

M=25

SIM

CFD

Figure 33. CA and CN at M∞=25 and α=30 deg. ORV-SB

concept. CFD and HPM comparison.

13


As shown, also in those cases it is confirmed that

HPM provides reliable results in the framework of a

phase-A design level.

5 Concluding Remarks

This paper provides a summary overview of

aerodynamic performances of three winged

vehicle concepts allowing to perform a number

of experiments on critical re-entry technologies.

Therefore, a possible re-entry mission scenario

has been provided and considered for the

appraisal of the concepts flight performance.

Results of forty seven fully three-dimensional

CFD computations of the flowfield past the

flying test beds have been presented for the

whole continuum flight regime. An assessment

of the vehicle aerodynamics performed through

low-order method design approach like

hypersonic panel methods are shown as well.

Real gas effects on the vehicles aerodynamics

have been discussed. Results show that these

effects for the ORV-WSB concept flying at

M∞=20 and for α ranging from 10 to 40 deg

seem to be negligible for lift and drag due to the

rather high streamlined vehicle configuration.

So real gas effect are expected to not affect

concept re-entry trajectory; while a significant

effect on vehicle pitching moment is expected

during descent varying from 8 to 50%.

Finally, aerodynamic analysis provided in this

work are considered applicable for the

prosecution of the trade-off design analysis of

the most promising aeroshape to choose for the

flying test bed.

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